Spacecraft attitude determination systems often include more than one attitude sensor. Some examples of typical attitude determination sensors are: 3-axes gyroscopes, star trackers, earth sensors, sun sensors, and beacon sensors. A multi-sensor based attitude determination system uses two or more attitude sensors to construct the spacecraft attitude. Typically, each individual attitude sensor generates attitude data at a moment in time different from the other attitude sensors. This is due to the asynchronous nature of the individual sensors, or to differences between sensor processing delays or data transmission delays. Accordingly, the sensor measurements available to the spacecraft control processor for controlling spacecraft attitude at a particular time, have different associated time moments. If the time mismatch is constant, the attitude error due to time mismatch can be accounted for, and calibrated out of the system. Thus, for example, in an orbit-normal steered spacecraft in a circular orbit, time-mismatched attitude sensor data will result in a constant attitude error which can be calibrated out of the attitude calculations.
The time discrepancy between the attitude measurement sensors can cause significant problems, however, when the spacecraft is experiencing dynamic motions such as agile slew, sun-nadir steering, yaw-flip, and spacecraft orbiting in a highly elliptical orbit (HEO). Time-mismatch of attitude sensor measurements is also a concern for traditional orbit-normal steered spacecraft in a circular orbit, if the time-mismatch is time varying. This will often be the case if the attitude sensors are running on their own clocks, independent from the system clock, and communicating with the spacecraft control processor through an asynchronous data bus.
The issue of time-mismatch between attitude measurement sensors is best illustrated by example. In a satellite having a stellar attitude sensor such as a star tracker and an inertial attitude sensor such a gyroscope, the stellar-inertial attitude determination system has two solutions--the inertial solution and the stellar solution. The inertial solution is obtained by integrating the gyroscope delta angle output, and the stellar solution is obtained by algebraic transformations of the star tracker measurements. Typically, these two solutions are not synchronized, and at any moment, the most recent solutions available to the spacecraft control processor do not match in time. Thus, a time-mismatch exists between the inertially derived attitude integrated from the gyroscope delta angles, and the attitude derived from the star tracker measurements. Thus, for example, if the spacecraft is stewing at 0.3 deg/sec and there is a 0.125 second mismatch between the stellar and inertial measurements, a discrepancy of 135 arcseconds exists. This can lead to inaccurate corrections of the integrated attitude solution. For spacecraft systems which require a few arcsecond or better attitude accuracy, this is not acceptable.